Combustor assembly for a turbine engine

ABSTRACT

A combustor assembly for a gas turbine engine includes a dome having a forward surface and an inner surface. The forward surface and the inner surface of the dome at least partially define a slot. The combustor assembly also includes a liner at least partially defining a combustion chamber and extending between an aft end and a forward end. The forward end of the liner is positioned within the slot of the dome. The forward end of the liner includes an axial interface surface and a radial interface surface. The axial interface surface defines a radial gap with the inner surface of the dome and the radial interface surface defines an axial gap with the forward surface of the dome. At least one of the radial gap or the axial gap is less than about 0.150 inches during operating conditions of the combustor assembly to prevent an undesirable airflow.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.16/654,571 filed on Oct. 16, 2019, which is a divisional of U.S. patentapplication Ser. No. 15/239,888 filed on Aug. 18, 2016, the entirecontents of each of which are hereby incorporated by reference in theirentireties.

FIELD OF THE INVENTION

The present subject matter relates generally to a gas turbine engine, ormore particularly to a combustor assembly for a gas turbine engine.

BACKGROUND OF THE INVENTION

A gas turbine engine generally includes a fan and a core arranged inflow communication with one another. Additionally, the core of the gasturbine engine general includes, in serial flow order, a compressorsection, a combustion section, a turbine section, and an exhaustsection. In operation, air is provided from the fan to an inlet of thecompressor section where one or more axial compressors progressivelycompress the air until it reaches the combustion section. Fuel is mixedwith the compressed air and burned within the combustion section toprovide combustion gases. The combustion gases are routed from thecombustion section to the turbine section. The flow of combustion gassesthrough the turbine section drives the turbine section and is thenrouted through the exhaust section, e.g., to atmosphere.

More commonly, non-traditional high temperature materials, such asceramic matrix composite (CMC) materials, are being used as structuralcomponents within gas turbine engines. For example, given an ability forCMC materials to withstand relatively extreme temperatures, there isparticular interest in replacing components within the combustionsection of the gas turbine engine with CMC materials. More particularly,an inner liner and an outer liner of gas turbine engines are morecommonly being formed of CMC materials.

By contrast, a dome within the combustion section may be formed of ametal material, with the inner and outer liners attached thereto. It canbe difficult, however, to provide cooling air to the inner and outerliners proximate the attachment locations of the inner and outer linersto the dome. Accordingly, a gas turbine engine, or more particularly, acombustor assembly of a gas turbine engine, capable of more effectivelyproviding a desired flow of cooling air to the inner and outer linersproximate the attachment locations of the inner and outer liners to thedome would be useful.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, a combustorassembly is provided for a gas turbine engine defining an axialdirection and a radial direction. The combustor assembly includes a domehaving a forward surface and an inner surface. The forward surface andthe inner surface at least partially define a slot. The combustorassembly also includes a liner at least partially defining a combustionchamber and extending between an aft end and a forward end. The forwardend of the liner is received within the slot of the dome. The forwardend of the liner includes an axial interface surface and a radialinterface surface. The axial interface surface defines a radial gap withthe inner surface of the dome and the radial interface surface definesan axial gap with the forward surface of the dome. At least one of theradial gap or axial gap is less than about 0.150 inches during operatingconditions of the combustor assembly.

In an exemplary aspect of the present disclosure, a method is providedfor manufacturing a combustor assembly of a gas turbine engine. Thecombustor assembly includes a liner and a dome. The dome includes aforward surface and an inner surface. The method includes forming aliner of a ceramic matrix composite material to include a baselinegeometry, and removing material from the liner to change the baselinegeometry to include an interface surface. The method also includesmounting the liner to the dome.

In another exemplary aspect of the present disclosure, a method forcooling a combustor assembly of a gas turbine engine is provided. Thecombustor assembly includes a liner and a dome, the dome including aforward surface and an inner surface at least partially defining a slot.The liner includes a forward end received within the slot. The methodincludes providing a cooling airflow to the slot defined by the forwardsurface and the inner surface of the dome. The method also includesproviding the cooling airflow through an axial gap defined between theforward end of the liner and the forward surface of the dome, the axialgap being less than about 0.150 inches. The method also includesproviding the cooling airflow through a radial gap defined between theforward end of the liner and the inner surface of the dome, the radialgap being less than about 0.020 inches. The method also includesproviding the cooling airflow to a combustion chamber defined at leastin part by the liner and the dome.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic cross-sectional view of an exemplary gas turbineengine according to various embodiments of the present subject matter.

FIG. 2 is a schematic, cross-sectional view of a combustor assembly inaccordance with an exemplary embodiment of the present disclosure.

FIG. 3 is a close up, cross-sectional view of an attachment point of theexemplary combustor assembly of FIG. 2 , where a forward end of an outerliner in accordance with an exemplary embodiment of the presentdisclosure is attached to an outer dome section.

FIG. 4 is a close-up, isolated view of the forward end of the exemplaryouter liner of FIG. 3 .

FIG. 5 is a schematic, axial view of the forward end of the exemplaryouter liner of FIG. 3 .

FIG. 6 side, cross-sectional view of an outer liner in accordance withan exemplary embodiment of the present disclosure.

FIG. 7 is a flow diagram of a method for manufacturing a combustorassembly of a gas turbine engine in accordance with an exemplary aspectof the present disclosure.

FIG. 8 is a flow diagram of a method for cooling a component of acombustor assembly of a gas turbine engine in accordance with anexemplary aspect of the present disclosure.

DETAILED DESCRIPTION OF THE INVENTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first”, “second”, and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms “forward”and “aft” refer to relative positions within a gas turbine engine, withforward referring to a position closer to an engine inlet and aftreferring to a position closer to an engine nozzle or exhaust. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows, and “downstream” refers to thedirection to which the fluid flows.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1 , the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1 , the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal centerline 12 provided forreference), a radial direction R, and a circumferential direction (i.e.,a direction extending about the axial direction A; not depicted). Ingeneral, the turbofan 10 includes a fan section 14 and a core turbineengine 16 disposed downstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22.

For the embodiment depicted, the fan section 14 includes a variablepitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 ina spaced apart manner. As depicted, the fan blades 40 extend outwardlyfrom disk 42 generally along the radial direction R. Each fan blade 40is rotatable relative to the disk 42 about a pitch axis P by virtue ofthe fan blades 40 being operatively coupled to a suitable actuationmember 44 configured to collectively vary the pitch of the fan blades 40in unison. The fan blades 40, disk 42, and actuation member 44 aretogether rotatable about the longitudinal axis 12 by LP shaft 36 acrossa power gear box 46. The power gear box 46 includes a plurality of gearsfor stepping down the rotational speed of the LP shaft 36 to a moreefficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 1 , the disk 42 iscovered by rotatable front nacelle 48 aerodynamically contoured topromote an airflow through the plurality of fan blades 40. Additionally,the exemplary fan section 14 includes an annular fan casing or outernacelle 50 that circumferentially surrounds the fan 38 and/or at least aportion of the core turbine engine 16. It should be appreciated that thenacelle 50 may be configured to be supported relative to the coreturbine engine 16 by a plurality of circumferentially-spaced outletguide vanes 52. Moreover, a downstream section 54 of the nacelle 50 mayextend over an outer portion of the core turbine engine 16 so as todefine a bypass airflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58 entersthe turbofan 10 through an associated inlet 60 of the nacelle 50 and/orfan section 14. As the volume of air 58 passes across the fan blades 40,a first portion of the air 58 as indicated by arrows 62 is directed orrouted into the bypass airflow passage 56 and a second portion of theair 58 as indicated by arrow 64 is directed or routed into the LPcompressor 22. The ratio between the first portion of air 62 and thesecond portion of air 64 is commonly known as a bypass ratio. Thepressure of the second portion of air 64 is then increased as it isrouted through the high pressure (HP) compressor 24 and into thecombustion section 26, where it is mixed with fuel and burned to providecombustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

It should be appreciated, however, that the exemplary turbofan engine 10depicted in FIG. 1 is by way of example only, and that in otherexemplary embodiments, the turbofan engine 10 may have any othersuitable configuration.

Referring now to FIG. 2 , a close-up cross-sectional view is provided ofa combustor assembly 100 in accordance with an exemplary embodiment ofthe present disclosure. For example, the combustor assembly 100 of FIG.2 may be positioned in the combustion section 26 of the exemplaryturbofan engine 10 of FIG. 1 . More particularly, FIG. 2 provides aside, cross-sectional view of the exemplary combustor assembly 100 ofFIG. 2 .

As shown, the combustor assembly 100 generally includes an inner liner102 extending between an aft end 104 and a forward end 106 generallyalong the axial direction A, as well as an outer liner 108 alsoextending between an aft end 110 and a forward end 112 generally alongthe axial direction A. The inner and outer liners 102, 108 together atleast partially define a combustion chamber 114 therebetween. The innerand outer liners 102, 108 are each attached to an annular dome. Moreparticularly, the annular dome includes an inner dome section 116attached to the forward end 106 of the inner liner 102 and an outer domesection 118 attached to the forward end 112 of the outer liner 108. Theinner and outer dome section 116, 118 may be formed integrally (oralternatively may be formed of a plurality of components attached in anysuitable manner) and may each extend along the circumferential directionC to define an annular shape. As will be discussed in greater detailbelow with reference to FIG. 3 , the inner and outer dome sections 116,118 each also include a forward surface 120 and an inner surface 121(i.e., inner relative to the combustion chamber 114) at least partiallydefining a slot 122 for receipt of the forward end 106 of the innerliner 102, and the forward end 112 of the outer liner 108, respectively.

The combustor assembly 100 further includes a plurality of fuel airmixers 124 spaced along a circumferential direction C and positioned atleast partially within the annular dome. More particularly, theplurality of fuel air mixers 124 are disposed at least partially betweenthe outer dome section 118 and the inner dome section 116 along theradial direction R. Compressed air from the compressor section of theturbofan engine 10 flows into or through the fuel air mixers 124, wherethe compressed air is mixed with fuel and ignited to create thecombustion gases 66 within the combustion chamber 114. The inner andouter dome sections 116, 118 are configured to assist in providing sucha flow of compressed air from the compressor section into or through thefuel air mixers 124. For example, the outer dome section 118 includes anouter cowl 126 at a forward end 128 and the inner dome section 116similarly includes an inner cowl 130 at a forward end 132. The outercowl 126 and inner cowl 130 may assist in directing the flow ofcompressed air from the compressor section 26 into or through one ormore of the fuel air mixers 124.

Moreover, the inner and outer dome sections 116, 118 each includeattachment portions configured to assist in mounting the combustorassembly 100 within the turbofan engine 10. For example, the outer domesection 118 includes an attachment extension 134 configured to bemounted to an outer combustor casing 136 and the inner dome section 116includes a similar attachment extension 138 configured to attach to anannular support member 140 within the turbofan engine 10. In certainexemplary embodiments, the inner dome section 116 may be formedintegrally as a single annular component, and similarly, the outer domesection 118 may also be formed integrally as a single annular component.It should be appreciated, however, that in other exemplary embodiments,the inner dome section 116 and/or the outer dome section 118 mayalternatively be formed by one or more components being joined in anysuitable manner. For example, with reference to the outer dome section118, in certain exemplary embodiments, the outer cowl 126 may be formedseparately from the outer dome section 118 and attached to the forwardend 128 of the outer dome section 118 using, e.g., a welding process.Similarly, the attachment extension 134 may also be formed separatelyfrom the outer dome section 118 and attached to the forward end 128 ofthe outer dome section 118 using, e.g., a welding process. Additionally,or alternatively, the inner dome section 116 may have a similarconfiguration.

Referring still to FIG. 2 , the exemplary combustor assembly 100 furtherincludes a heat shield 142 positioned around the fuel air mixer 124depicted. The exemplary heat shield 142, for the embodiment depicted, isattached to and extends between the outer dome section 118 and the innerdome section 116. The heat shield 142 is configured to protect certaincomponents of the turbofan engine 10 from the relatively extremetemperatures of the combustion chamber 114.

For the embodiment depicted, the inner liner 102 and the outer liner 108are each formed of a ceramic matrix composite (CMC) material, which is anon-metallic material having high temperature capability. Exemplary CMCmaterials utilized for such liners 102, 108 may include silicon carbide,silicon, silica or alumina matrix materials and combinations thereof.Ceramic fibers may be embedded within the matrix, such as oxidationstable reinforcing fibers including monofilaments like sapphire andsilicon carbide (e.g., Textron's SCS-6), as well as rovings and yarnincluding silicon carbide (e.g., Nippon Carbon's NICALON®, UbeIndustries' TYRANNO®, and Dow Corning's SYLRAMIC®), alumina silicates(e.g., Nextel's 440 and 480), and chopped whiskers and fibers (e.g.,Nextel's 440 and SAFFIL®), and optionally ceramic particles (e.g.,oxides of Si, Al, Zr, Y and combinations thereof) and inorganic fillers(e.g., pyrophyllite, wollastonite, mica, talc, kyanite andmontmorillonite). CMC materials may have coefficients of thermalexpansion in the range of about 1.3×10⁻⁶ in/in/° F. to about 3.5×10⁻⁶in/in/° F. in a temperature of approximately 1000-1200° F.

By contrast, the annular dome, including the inner dome section 116 andouter dome section 118, may be formed of a metal, such as a nickel-basedsuperalloy (having a coefficient of thermal expansion of about8.3-8.5×10⁻⁶ in/in/° F. in a temperature of approximately 1000-1200° F.)or cobalt-based superalloy (having a coefficient of thermal expansion ofabout 7.8-8.1×10⁻⁶ in/in/° F. in a temperature of approximately1000-1200° F.).

Referring still to FIG. 2 , at the aft end 104 of the inner liner 102and at the aft end 110 of the outer liner 108, the combustor assembly100 includes an inner piston ring seal 146 and an outer piston ring seal148, respectively. The inner piston ring seal 146 is attached to aninner piston ring holder 150 extending from and attached to an interiorcasing (which for the embodiment depicted is the annular support member140). Similarly, the outer piston ring seal 148 is attached to an outerpiston ring holder 152 extending from and attached to an outer casing(which for the embodiment depicted includes the outer combustor casing136 and an outer turbine casing 154). The inner piston ring holder 150and the outer piston ring holder 152 are configured to accommodate anexpansion of the inner liner 102 and the outer liner 108 generally alongthe axial direction A, as well as generally along the radial directionR.

Referring still to FIG. 2 , and as is discussed above, the combustiongases 66 flow from the combustion chamber 114 into and through theturbine section of the turbofan engine 10 where a portion of thermaland/or kinetic energy from the combustion gases 66 is extracted viasequential stages of turbine stator vanes and turbine rotor blades. Astage one (1) turbine blade 156 is depicted schematically in FIG. 2 ,aft of the combustor assembly 100.

Referring now to FIG. 3 , a close up, schematic, cross-sectional view isdepicted of an attachment point where the forward end 112 of the outerliner 108 is mounted to the outer dome section 118 within the slot 122of the outer dome section 118. The view of FIG. 3 (and also of FIG. 2 )is during operating conditions of the combustor assembly. Morespecifically, FIG. 3 depicts the combustor assembly during conditionswhere fuel is being burned within the combustion chamber 114 to generatecombustion gasses and all components of the combustor assembly 100 havereached a steady temperature.

To allow for a relative thermal expansion between the outer liner 108and the outer dome section 118, as well as between the inner liner 102and the inner dome section 116, a plurality of mounting assemblies 144are used to attach the outer liner 108 to the outer dome section 118 andthe inner liner 102 to the inner dome section 116. More particularly,the mounting assemblies 144 attach the forward end 112 of the outerliner 108 to the outer dome section 118 within the slot 122 of the outerdome section 118 and the forward end 106 of the inner liner 102 to theinner dome section 116 within the slot 122 of the inner dome section 116(see FIG. 2 ). As stated, the slots 122 are at least partially definedby the forward surfaces 120 and inner surfaces 121 of the respectiveinner and outer dome sections 116, 118, and further, the slots 122receive the forward ends 106, 112 of the inner and outer liners 102,108, respectively.

Referring particularly to the forward end 112 of the outer liner 108 andthe outer dome section 118 depicted in FIG. 3 , the outer dome section118 includes a base plate 158 and a yoke 160. The base plate 158 and theyoke 160 each extend substantially parallel to one another, which forthe embodiment depicted is a direction substantially parallel to theaxial direction A of the turbofan engine 10 (see also FIG. 2 ). Notably,the slot 122 is defined between the base plate 158 and the yoke 160.Further, in certain exemplary embodiments, the yoke 160 may extendcircumferentially with the outer dome section 118, tracking the baseplate 158. With such a configuration, the slot 122 may be considered anannular slot. However, in other embodiments, the yoke 160 may include aplurality of circumferentially spaced tabs, each of the individual tabsof the yoke 160 defining individual segmented portions of the slot 122with the base plate 158.

The exemplary mounting assembly 144 depicted extends through the yoke160 of the outer dome section 118, the forward end 112 of the outerliner 108 (positioned in the slot 122), and the base plate 158 of theouter dome section 118. More particularly, for the embodiment depicted,the mounting assembly 144 includes a pin 162 and a bushing 164. The pin162 includes a head 166 and a shank 168, the shank 168 extending throughthe yoke 160, the forward end 112 of the outer liner 108 (positioned inthe slot 122), and the base plate 158. A nut 170 is attached to a distalend of the shank 168 of the pin 162. In certain exemplary embodiments,the pin 162 may be configured as a bolt and the nut 170 may be rotatablyengaged with a threaded portion of the pin 162 (at, e.g., the distal endof the shank 168) for tightening the mounting assembly 144.Alternatively, however, in other exemplary embodiments the pin 162 andnut 170 may have any other suitable configuration. For example, in otherexemplary embodiments, the pin 162 may include a shank 168 defining asubstantially smooth cylindrical shape and the nut 170 may be configuredas a clip.

Additionally, the bushing 164 is generally cylindrical in shape andpositioned around the shank 168 of the pin 162 within the slot 122. Forthe embodiment depicted, the bushing 164 is pressed between the yoke 160and the base plate 158 by tightening the nut 170 on the pin 162.Moreover, for the embodiment depicted, the mounting assembly 144includes a metal grommet 172 positioned around the bushing 164 and pin162. The grommet 172 is positioned in an opening 174 in the forward end112 of the outer liner 108. The grommet 172 includes an outer collar 176positioned adjacent to an outside surface 178 of the outer liner 108 andan inner collar 180 positioned adjacent to an inside surface 182 of theouter liner 108. The grommet 172 additionally includes a body 184. Themetal grommet 172 may reduce an amount of wear on the forward end 112 ofthe outer liner 108 as the outer liner 108 moves inwardly and outwardlygenerally along the radial direction R relative to the outer domesection 118.

It should be appreciated, however, that although the forward end 112 ofthe outer liner 108 is attached to the outer dome section 118 using theexemplary attachment assembly 144 depicted and described herein, inother embodiments of the present disclosure, the attachment assembly 144may have any other suitable configuration, and further still in otherembodiments, any other suitable attachment assembly may be used.

Referring still to FIG. 3 , the forward end 112 of the outer liner 108depicted further includes an axial interface surface 186 and a radialinterface surface 188. The axial interface surface 186 is configured asa portion of the forward end 112 of the outer liner 108 facing the baseplate 158 of the outer dome section 118, or more particularly, facingthe inner surface 120 of the outer dome section 118. The radialinterface surface 188 is configured as a portion of the forward end 112of the inner liner facing the forward surface 121 of the outer domesection 118. For the embodiment depicted, the axial interface surface186 and inner surface 120 each extend in a direction parallel to theaxial direction A, and the radial interface surface 188 and forwardsurface 121 each extend in a direction parallel to the radial directionR.

Moreover, for the embodiment depicted, the axial interface surface 186defines a radial gap 190 with the inner surface 120 of the outer domesection 118 and the radial interface surface 188 defines an axial gap192 with the forward surface 121 of the outer dome section 118. For theembodiment depicted, at least one of the radial gap 190 or axial gap 192is less than about 0.150 inches during operating conditions of thecombustor assembly 100. More particularly, for the embodiment depicted,at least one of the radial gap 190 or axial gap 192 is less than about0.020 inches during operating conditions of the combustor assembly 100.

For example, referring still to the embodiment of FIG. 3 , the radialgap 190 is less than about 0.020 inches during operating conditions ofthe combustor assembly 100 and the axial gap 192 is less than about0.150 inches during operating conditions of the combustor assembly 100.More specifically, for the embodiment depicted, the radial gap 190defined by the axial interface surface 186 of outer liner 112 with theinner surface 120 of the outer dome section 118 is between about 0.005inches and about 0.015 inches, and the axial gap 192 defined by theradial interface surface 188 of the outer liner 112 with the forwardsurface 121 of the outer dome section 118 is between about 0.050 inchesand about 0.100 inches. It should be appreciated, that as used herein,terms of approximation, such as “about” and “approximately,” refer tobeing within a ten percent (10%) margin of error.

The combustor assembly 100 may be designed such the radial and axialgaps 190, 192 defined by the axial interface surface 186 with the innersurface 120 and by the radial interface surface 188 with the forwardsurface 121 allow for only a predetermined amount of airflowtherethrough into the combustion chamber 114. Notably, allowing such aflow of air during operating conditions of the combustor assembly 100may ensure relatively hot combustion gases within the combustion chamber114 do not flow into and/or through the slot 122 of the outer domesection 118, potentially damaging certain components of the combustorassembly 100.

Referring now to FIG. 4 , a close-up, isolated view of the forward end112 of the outer liner 108 of FIGS. 2 and 3 is provided. Morespecifically, FIG. 4 shows the forward end 112 of the outer liner 108including a baseline geometry. The outer liner 108 may be formed, e.g.,of a CMC material, such that the outer liner 108 includes the baselinegeometry. Subsequently, the outer liner 108 may be machined to definethe axial interface surface 186 and the radial interface surface 188.The baseline geometry is depicted in phantom lines in FIG. 4 .Accordingly, for the embodiment depicted, the axial interface surface186 is a machined surface and a radial interface surface 188 is also amachined surface. By forming the forward end 112 of the outer liner 108to include a baseline geometry and subsequently machining down thebaseline geometry such that the axial interface surface 186 and theradial interface surface 188 are defined, the forward end 112 of theouter liner 108 may define the desired radial and axial gaps 190, 192with the inner surface 120 and forward surface 121, respectively, of theouter dome section 118 during operating conditions of the combustorassembly 100 once the forward end 112 of the outer liner 108 is receivedwithin the slot 122 of the outer dome section 118.

Furthermore, referring now to FIG. 5 , an axial view of the forward end112 of the outer liner 108 of FIGS. 2 and 3 is provided. Morespecifically, FIG. 5 also shows the forward end 112 of the outer liner108 including the baseline geometry (shown in phantom). As is depicted,certain manufacturing methods of CMC components may make it difficult toform the forward end 112 of the outer liner 108 to include a preciselycircular inner surface relative to a centerline 12 of the gas turbineengine. Accordingly, by forming the forward end 112 of the outer liner108 to include the baseline geometry (depicted in phantom), andsubsequently machining down the baseline geometry of the forward end112, such that the axial interface surface 186 and radial interfacesurface 188 (see FIG. 4 ) are defined may ensure more consistent radialand axial gaps 190, 192 are defined with the inner surface 120 and theforward surface 121, respectively, of the outer dome section 118 alongthe circumferential direction C. More specifically, machining down thebaseline geometry of the forward end 112 of the outer liner 108 mayensure the axial interface surface 186 defines a more precise circularshape with respect to the centerline 12 of the gas turbine engine, andtherefore a more consistent radial gap 190 with the inner surface 120 ofthe outer dome section 118.

Moreover, referring back to FIG. 2 , it should be appreciated that theforward end 106 of the inner liner 102 may be formed in substantiallythe same manner as the forward end 112 of the outer liner 108, and alsothat the forward end 106 of the inner liner 102 may be attached to theinner dome section 116 in substantially the same manner that the forwardend 112 of the outer liner 108 is attached to the outer dome section118. For example, the mounting assemblies 144 attaching the forward end106 of the inner liner 102 to the inner dome section 116 may beconfigured in substantially the same manner as the mounting assemblies144 attaching the forward end 112 of the outer liner 108 to the outerdome section 118. Additionally, the inner dome section 116 may define aforward surface 121 and an inner surface 120 (i.e., inner relative tothe combustion chamber 114). Additionally, the forward end 106 of theinner liner 102 may include an axial interface surface 186 facing theinner surface 120 of the inner dome section 116 and a radial interfacesurface 188 facing the forward surface 121 of the inner dome section116. The axial interface surface 186 of the forward end 106 of the innerliner 102 may define a radial gap with the inner surface 120 of theinner dome section 116 less than about 0.020 inches during operatingconditions the combustor assembly 100, and further the radial interfacesurface 188 of the forward end 106 of the inner liner 102 may define anaxial gap with the forward surface 121 of the inner dome section 116less than about 0.150 inches during operating conditions of thecombustor assembly 100.

Further, still, it should be appreciated that in other exemplaryembodiments, other portions of the liners, and/or other components of agas turbine engine may be formed in a similar manner to ensure suchportions of the liners (or other component of a gas turbine engine)define a desired thickness and/or clearance with adjacent components.For example, referring now to FIG. 6 , a side, cross-sectional view isprovided of a liner in accordance with an exemplary embodiment of thepresent disclosure. In certain exemplary embodiments the liner may be anouter liner 108 configured in substantially the same manner as the outerliner 108 described above with reference to FIGS. 2 through 5 .Accordingly, the same numbers may refer to same or similar part.

For example, as is depicted, the outer liner 108 of FIG. 6 extendsgenerally between an aft end 110 and a forward end 112. The outer liner108 additionally defines a midspan region 193 positioned between theforward end 112 and the aft end 110. Moreover, the outer liner 108generally defines an inner surface 194, defining at least in part acombustion chamber when installed (see FIG. 2 ), and an opposite, outersurface 195. As with the outer liner 108 described above with referenceto FIGS. 4 and 5 , the outer liner 108 of FIG. 6 is formed, e.g., of aCMC material, such that the outer liner 108 includes a baselinegeometry. Subsequent to formation, the outer liner 108 may be machinedin various locations to define certain interface surfaces as may bedesired for the specific application. More particularly, as is depictedvia the various machine lines 196 in phantom, the forward end 112 of theouter liner 108 is machined to define an axial interface surface 197,the midspan region 193 of the outer liner 108 is machined to define amidspan interface surface 198 at the outer surface 195, and the aft end110 of the outer liner 108 is machined to define an aft end interfacesurface 199. Notably, the exemplary outer liner 108 of FIG. 6 is formedsuch that the remaining portion of the outer liner 108 (i.e., aftermachining) includes a sufficient mass or thickness to maintain anydesired mechanical properties (e.g., stiffness, strength, flexibility,etc.).

It should be appreciated, however, that in still other exemplaryembodiments, any other suitable portions of the baseline geometry of theouter liner 108 of FIG. 6 may be machined to define a desired interfacesurface. Moreover, in still other embodiments, any other suitablecomponent of a gas turbine engine, and specifically any other suitablecomponent formed of a CMC material, may be configured in a similarmanner as exemplary outer liner 108 of FIG. 6 to include a baselinegeometry that may be machined down to define a desired interfacesurface.

Referring now to FIG. 7 , a method (200) for manufacturing a combustorassembly of a gas turbine engine in accordance with an exemplary aspectof the present disclosure is provided. The exemplary method (200)depicted in FIG. 7 may be used to manufacture the exemplary combustorassembly described above with reference to FIGS. 2 through 5 .Accordingly, the exemplary combustor assembly manufactured according tothe exemplary method (200) may include a liner and a dome, with the domeincluding a forward surface and an inner surface, the forward surfaceand the inner surface together at least partially defining a slot.

The exemplary method (200) includes at (202) forming a liner to includea forward end having a baseline geometry. In certain embodiments, theliner may be an inner liner of the combustor assembly, or alternatively,may be an outer liner of the combustor assembly. Notably, for theembodiment depicted, forming the liner at (202) includes at (204)forming the liner of a ceramic matrix composite material.

Further, the exemplary method (200) includes at (206) removing materialfrom the forward end of the liner to change the baseline geometry toinclude an axial interface surface and a radial interface surface. Incertain exemplary embodiments, removing material from the forward end ofthe liner at (206) includes at (208) machining the forward end of theliner to define the axial interface surface, and at (210) machining theforward end of the liner to define the radial interface surface.

Moreover, referring still to FIG. 7 , the exemplary method (200)additionally includes at (212) mounting the forward end of the liner atleast partially within the slot of the dome, such that the axialinterface surface defines a radial gap with the inner surface of thedome and the radial interface surface defines an axial gap with theforward surface the dome. For the exemplary aspect depicted, at leastone of the axial gap or the radial gap is less than about 0.150 inchesduring operating conditions of the combustor assembly, or moreparticularly, less than about 0.020 inches during operating conditionsof the combustor assembly. Specifically, in at least certain exemplaryaspects, the radial gap may be less than about 0.020 inches duringoperating conditions of the combustor assembly and the axial gap may beless than about 0.150 inches during operating conditions of thecombustor assembly.

A combustor assembly manufactured in accordance with one or moreexemplary aspects of the present disclosure may ensure that a desiredamount of airflow is provided through the gaps defined between theforward end of the liner and the dome during operating conditions of thecombustor assembly, such that relatively hot combustion gases do notflow through the slot of the combustor dome, potentially damagingcertain components of the combustor assembly.

Furthermore, referring now to FIG. 8 , a method (300) for cooling acombustor assembly of a gas turbine engine in accordance with anexemplary aspect of the present disclosure is provided. The exemplarymethod (300) depicted in FIG. 8 may be used to cool the exemplarycombustor assembly described above with reference to FIGS. 2 through 5 .Accordingly, the exemplary combustor assembly of the exemplary method(300) may include a liner and a dome, with the dome including a forwardsurface and an inner surface, the forward surface and the inner surfacetogether at least partially defining a slot. Further, the liner mayinclude a forward end received within the slot of the dome.

The exemplary method (300) includes at (302) providing a cooling airflowto the slot defined by the forward surface and the inner surface of thedome. The cooling airflow may be a portion of an airflow through acompressor section of the gas turbine engine. For example, providing thecooling airflow to the slot at (302) may include providing a portion ofan airflow through the compressor section over a forward surface of thedome to the slot.

Additionally, the exemplary method (300) includes at (304) providing thecooling airflow through an axial gap defined between the forward end ofthe liner and the forward surface of the dome, the axial gap being lessthan about 0.150 inches. The exemplary method (300) additionallyincludes at (306) providing the cooling airflow through a radial gapdefined between the forward end of the liner and the inner surface ofthe dome, the radial gap being less than about 0.020 inches. Morespecifically, the axial gap may be defined between a radial interfacesurface of the forward end of the liner and the forward surface of thedome, and further, the radial gap may be defined between an axialinterface surface of the forward end of the liner and the inner surfaceof the dome.

Moreover, as is also depicted, the exemplary method (300) includes at(308) providing the cooling airflow to a combustion chamber defined atleast in part by the liner and the dome. Cooling a combustor assembly inaccordance with the exemplary method (300) may ensure a sufficientamount of cooling airflow is provided through the slot, around theforward end of the liner, and to the combustion chamber to preventcombustion gases from flowing back through the slot. Moreover, byproviding the cooling airflow through the radial and axial gaps,excessive cooling airflow may be prevented from flowing therethrough.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A combustor assembly for a gas turbine enginedefining a centerline, an axial direction, and a radial direction, thecombustor assembly comprising: a dome defining a slot; a liner formed ofa ceramic matric composite material and at least partially defining acombustion chamber, the liner extending between an aft end and a forwardend, the forward end of the liner configured to be received within theslot of the dome, the forward end comprising an opening, a radialinterface surface, and an axial interface surface; an axial gap locatedbetween a forward surface of the dome and the radial interface surface;a radial gap located between an inner surface of the dome and the axialinterface surface; and at least one mounting assembly configured tosecure the liner to the dome, the at least one mounting assemblycomprising: a grommet located in the opening in the liner, the grommetconfigured to reduce wear on the forward end of the liner; a bushinginstalled in the opening of the liner and within the grommet; a collaradjacent an outer surface of the liner; a pin inserted through thebushing and extending through the dome and the opening in the liner; anda nut adjacent the inner surface of the dome and configured to tightenon the pin.
 2. The combustor assembly of claim 1, wherein the grommet islocated between the liner and the bushing.
 3. The combustor assembly ofclaim 1, wherein the radial interface surface is a machined surface todefine the axial gap.
 4. The combustor assembly of claim 1, wherein theaxial interface surface is a machined surface to define the radial gap.5. The combustor assembly of claim 1, wherein the axial gap and theradial gap are configured to allow a predetermined amount of airflowtherethrough into the combustion chamber.
 6. The combustor assembly ofclaim 1, the dome further comprising a yoke and a base plate, whereinthe collar is an outer collar positioned between the outer surface ofthe liner and an inner surface of the yoke, the combustor assemblyfurther comprising an inner collar adjacent an inner surface of theliner, the inner collar positioned between the inner surface of theliner and the base plate.
 7. The combustor assembly of claim 1, whereinthe dome is formed of a metal material.
 8. The combustor assembly ofclaim 1, wherein the outer surface of the liner defines a cold sidesurface and a mid-span region, and wherein the liner defines an outerinterface surface on the cold side surface within the mid-span region.9. The combustor assembly of claim 1, wherein the liner is an outerliner and wherein the dome is an outer dome section.
 10. The combustorassembly of claim 1, wherein the liner is an inner liner and wherein thedome is an inner dome section.
 11. The combustor assembly of claim 1,the dome further comprising a yoke and a base plate, wherein the pinextends through the yoke, the forward end of the liner, and the baseplate.
 12. The combustor assembly of claim 1, wherein the axialinterface surface is defined at a local area of the liner having anincreased thickness relative to a first area immediately forward of theaxial interface surface and a second area immediately aft of the axialinterface surface.
 13. The combustor assembly of claim 1, wherein theaxial interface surface and the inner surface of the dome extend in adirection parallel to the axial direction.
 14. The combustor assembly ofclaim 1, wherein the radial interface surface and the forward surface ofthe dome extend in a direction parallel to the radial direction.
 15. Thecombustor assembly of claim 1, the dome comprising a yoke and a baseplate, the yoke and the base plate defining the slot therebetween. 16.The combustor assembly of claim 15, wherein the nut is configured topress the bushing between the yoke and the base plate.
 17. The combustorassembly of claim 1, wherein at least one of the axial gap or the radialgap is less than about 0.150 inches during operating conditions of thecombustor assembly.
 18. The combustor assembly of claim 17, wherein theradial gap is less than about 0.020 inches during the operatingconditions of the combustor assembly.
 19. The combustor assembly ofclaim 18, wherein the radial gap is between 0.005 and 0.015 inchesduring the operating conditions of the combustor assembly.
 20. Thecombustor assembly of claim 19, wherein the axial gap is less than about0.150 inches during the operating conditions of the combustor assembly.